In addition to their traditional propulsion functions, gas turbine engines are used as auxiliary power units aboard many types of aircraft to supply pneumatic power and/or shaft horsepower. A gas turbine auxiliary power unit includes in flow series arrangement a compressor, a bleed port, a combustor, and a turbine driving the compressor. A gearbox is drivingly coupled to the turbine and has a variety of aircraft accessories mounted thereto. Because the bleed port is disposed between the compressor and the combustor, these engines are referred to as integral bleed engines. When required, compressed air is bled through the bleed port and delivered to the aircraft where it is used in the environmental control system or for main engine starting. A load control valve is disposed in the bleed port to control the amount of bleed air. In an alternative configuration, a second compressor is operably coupled to the turbine and provides any required bleed flow. This type of configuration in referred to as a load compressor engine.
Regardless of the configuration, gas turbine auxiliary power units generally have two primary modes of operation. The first mode is maximum bleed flow in which the load control valve fully opens the bleed port and the maximum amount of compressed air is delivered to the aircraft. The second mode is maximum horsepower in which the control valve closes the bleed port and the horsepower generated by the engine is used to drive the accessories. As the engine transitions from the first mode to the second mode, the bleed port is closing which causes the pressure downstream of the compressor to rise. Should this pressure rise above the pumping capacity of the compressor, a surge will occur. During a surge, the direction of air flow in the compressor, reverses. This reversal of flow direction can be violent causing loud bangs and structural damage.
A conventional technique for preventing surge in these circumstances is to provide a surge valve within the bleed port. The surge valve is smaller than the load control valve and opens as the load control valve closes, thus limiting the rise of pressure downstream of the compressor. Unfortunately, the air bled through surge valve is dumped overboard. This dumped air is lost energy which must be compensated for by increasing the fuel flow to the combustor.
In order to eliminate the surge control valve and its associated losses, it has been proposed to employ a variable geometry diffuser within the compressor. It is well known in the art that variable geometry diffusers can improve an engine's power range and efficiency. With a variable geometry diffuser the engine can operate in the first mode with the diffuser in its maximum open setting and in the second mode with the diffuser in its minimum setting. However, even with a variable geometry diffuser the compressor can still be driven into a surge condition if the diffuser setting is not carefully adjusted as the engine transitions from the first mode to the second mode.
Accordingly, there is a need for a control and method that prevents compressor surge by adjusting the area of a variable geometry diffuser as the engine transitions from one operating condition to another.
Some aircraft manufactures desire to provide a signal to the auxiliary power unit when a change in the unit's operating conditions is required. For these aircraft a simple on/off the surge prevention control is provided that opens and closes the diffuser area in response to the aircraft signal. Additionally, the on/off control will close the diffuser in the event that it receives a faulty aircraft signal and a surge condition become imminent. Other manufacturers desire not to provide such signals. For these applications the surge prevention control must be self contained and able to modulate the diffuser area in response to a variety of measurements from sensors mounted in the gas turbine engine.